ALBERT

Description: A theoretical analysis of the radial temperature distribution through the rotor and constant cross sectional area blades near the coolant passages of liquid cooled gas turbines was made. The analysis was applied to obtain the rotor and blade temperatures of a specific turbine using a gas flow of 55 pounds per second, a coolant flow of 6.42 pounds per second, and an average coolant temperature of 200 degrees F. The effect of using kerosene, water, and ethylene glycol was determined. The effect of varying blade length and coolant passage lengths with water as the coolant was also determined. The effective gas temperature was varied from 2000 degrees to 5000 degrees F in each investigation.

Description: A theoretical analysis of the cross-sectional temperature distribution of a water-cooled turbine blade was made using the relaxation method to solve the differential equation derived from the analysis. The analysis was applied to specific turbine blade and the studies icluded investigations of the accuracy of simple methods to determine the temperature distribution along the mean line of the rear part of the blade, of the possible effect of varying the perimetric distribution of the hot gas-to -metal heat transfer coefficient, and of the effect of changing the thermal conductivity of the blade metal for a constant cross sectional area blade with two quarter inch diameter coolant passages.

Description: The performance at inlet pressure of 21 inches mercury absolute and inlet temperature of 538 R for the 10-stage axial-flow X24C-2 compressor from the X24C-2 turbojet engine was investigated. the peak adiabatic temperature-rise efficiency for a given speed generally occurred at values of pressure coefficient fairly close to 0.35.For this compressor, the efficiency data at various speeds could be correlated on two converging curves by the use of a polytropic loss factor derived.

Description: As part of an investigation to increase the power output of the V-1710-93 engine at altitude, the engine-stage supercharger was combined with a constant-area vaneless diffuser designed to improve the performance of the engine-stage supercharger at the rated engine operating point. The performance of the modified supercharger was investigated in a variable-component supercharger test rig and compared with that of the standard supercharger with an 8-vaned diffuser. A separate evaluation of the component efficiencies and a study of the flow characteristics of the modified supercharger was made possible by internal diffuser instrumentation. At the volume flow required by the engine for rated operating conditions, the modified supercharger increased the over-all adiabatic efficiency 0.05 and the over-all pressure coefficient 0.035. Furthermore, the capacity of the engine-stage supercharger was increased by replacing the standard 8-vaned diffuser with the vaneless diffuser. The peak over-all adiabatic efficiency for the modified supercharger, however, was 0.05 to 0.07 lower than that of the standard unit over the range of tip speeds investigated. The improved performance of the modified supercharger at rated engine operating conditions resulted from a shift of the point of peak adiabatic efficiency and pressure coefficient of the standard supercharger to a higher flow. The energy loss through the vaneless diffuser was found to be small. Because of the restricted diffuser diameter, however, diffusion was inadequate, which resulted in a relatively small static-pressure rise through the diffuser, high diffuser-exit velocities, and excessive collector-case losses.

Description: The present treatise reports on theoretical investigations and test-stand measurements which were carried out in the BMW Flugmotoren GMbH in developing the hollow blade for exhaust gas turbines. As an introduction the temperature variation and the stress on a turbine blade for a gas temperature of 900 degrees and circumferential velocities of 600 meters per second are discussed. The assumptions onthe heat transfer coefficients at the blade profile are supported by tests on an electrically heated blade model. The temperature distribution in the cross section of a blade Is thoroughly investigated and the temperature field determined for a special case. A method for calculation of the thermal stresses in turbine blades for a given temperature distribution is indicated. The effect of the heat radiation on the blade temperature also is dealt with. Test-stand experiments on turbine blades are evaluated, particularly with respect to temperature distribution in the cross section; maximum and minimum temperature in the cross section are ascertained. Finally, the application of the hollow blade for a stationary gas turbine is investigated. Starting from a setup for 550 C gas temperature the improvement of the thermal efficiency and the fuel consumption are considered as well as the increase of the useful power by use of high temperatures. The power required for blade cooling is taken into account.

Description: Design data are presented for the graphical construction of two-dimensional sharp-edge-throat supersonic nozzles of minimum length for test-section Mach numbers from 1.20 to 10.00. The method of characteristics used in the design is briefly reviewed.

Description: An investigation of the pressure distribution on the fuselage nose and the pilot canopy of a supersonic airplane model has been conducted at a Mach number of 1.90 over a wide range of angles of attack and yaw. Boundary layer separation apparently occurred from the upper surface at angles of attack above 24 degrees and from the lower surface at minus 15 degrees. No separation from the sides of the fuselage was evident at yaw angles up to 12 degrees.

Description: Computations were made to determine the temperature distribution and cooling of solid gas-turbine blades.A range of temperatures was used from 1500 degrees to 2500 degrees F, blade-root temperatures from 100 degrees to 1000 degrees F, blade thermal conductivity from 8 to 220 BTU/(hr)(sq ft)(degrees F/ft), and net gas to metal heat transfer coefficients from 75 to 250 BTU/(hr)(sq ft)(degrees F).

Description: Measurements on three tubes with flow regulated by suction at the trainling edge of the tube are described. It was possible to vary the mass of air flowing through the tube over a large range. Such tubes could be used for shrouded propellers.

Description: Effect of inlet-air pressure and temperature on the performance of the X24-2 10-Stage Axial-Flow Compressor from the X24C-2 turbojet engine was evaluated. Speeds of 80, 89, and 100 percent of equivalent design speed with inlet-air pressures of 6 and 12 inches of mercury absolute and inlet-air temperaures of approximately 538 degrees, 459 degrees,and 419 degrees R ( 79 degrees, 0 degrees, and minus 40 degrees F). Results were compared with prior investigations.

Description: A calulation of the flow in turbine blading is reported that includes the calculation of effect of centrifugal force. Frictional losses on the stator blades and rotor blades are allowed.

Description: An investigation was made in the Cleveland Altitude wind tunnel to determine the performance of an Aeroproducts H20C-162-X11M2 four-blade propeller on a YP-47M airplane at high blade loadings and high engine powers. The propeller characteristics were obtained for a range of power coefficients from 0.30 to 1.00 at free-stream Mach numbers of 0.40 and 0.50. The results of the force measurements are indicative only of trends in propeller efficiency with changes in power coefficient and advance-diameter ratio because unknown interference effects existed during the investigation. At a free-stream Mach number of 0.40, the envelopes of the efficiency curves decreased about 11% between advance-diameter ratios of 2.40 and 4.40. An increase in power coefficient from 0.30 to 0.80 at an advance-diameter ratio of 2.40 had little effect on the propeller efficiency. A change in power coefficient from 0.40 to 1.00 at an advance-diameter ratio of 4.40 increased the propeller efficiency by about 40%. For conditions below the stall the thrust loading on the outboard blade sections increased more rapidly than on the inboard sections as the power coefficient was increased or as the advance-diameter ratio was decreased. For conditions beyond the stall, the thrust loading decreased on the outboard sections and increased on the inboard sections.

Description: An investigation of the antiknock effectiveness of various additive-water solutions when used as internal coolants has been conducted at the NACA Cleveland laboratory. Nine compounds have been previously run in a CFR engine and the results are presented. In an effort to find a good anti-knock-coolant additive with more desirable physical properties than those of the nine compounds previously investigated, water solutions of four alkyl amines, three alkanolamines, six amides, and eight heterocyclic compounds were investigated and the results are presented.

Description: Qualitative investigations have shown that use of the NACA injection impeller with the R-3350 engine increases the inertia of the fuel-injection system and, when the standard fuel-metering system is used, this increase in inertia results in poor engine acceleration characteristics. This investigation was therefore undertaken to determine whether satisfactory acceleration characteristics of the engine equipped with the injection impeller could be obtained by simple modifications to the fuel-monitoring system. The engine was operated with two types of carburetor; namely, a hydraulic-metering carburetor incorporating a vacuum-operated accelerating pump and a direct-metering carburetor having a throttle-actuated accelerating pump. The vacuum-operated accelerating pump of the hydraulic-metering carburetor was modified to produce satisfactory accelerations by supplementing the standard air chamber with an additional 75-cubic spring. The throttle-actuated accelerating pump of the direct-metering carburetor was modified to produce satisfactory accelerations by replacing the standard 0.028-inch-diameter bleed in the load-compensator balance line with a smaller bleed of 0.0225-inch diameter. The results of this investigation indicated that both carburetors can be easily modified to produce satisfactory acceleration characteristics of the engine and no definite choice between the types of carburetor and accelerating pump can be made. Use of the direct-metering carburetor, however, probably resulted in better fuel distribution to the cylinders during the acceleration period and reduced the backfire hazard because all the fuel is introduced through the injection impeller.

Description: Three modifications of the auxiliary-stage supercharger for the V-1710-93 engine were designed and tested as part of an investigation to improve the power output and the altitude performance of the engine. A 12-vane diffuser was substituted for the standard 11-vane diffuser, and a vaneless discharge passage and a modified scroll were designed to increase the flow capacity of the supercharger and thereby to increase the performance at the high volume flows required by the engine. With the 12-vane diffuser installed and the carburetor replaced by an adapter, the equivalent volume flow at the peak efficiency point was increased 25 percent at the lowest speed investigated and 9.5 percent at the highest speed. When the carburetor was used, any increase in equivalent volume flow was masked by choking in the carburetor. A small decrease in the peak adiabatic efficiency resulted from using the 12-vane diffuser. At the high volume flows where the supercharger is required to operate, the performance was improved by the 12-vane diffuser. With the vaneless discharge passage, the surge-free range of the supercharger was increased 35 percent at the lowest tip speed investigated by increasing the maximum air flow. The maximum air flow at high tip speeds was again limited by choking in the carburetor, which masked the effect of the vaneless discharge passage on the maximum air flow. At the high volume flows near the operating point of the supercharger, the performance with the vaneless discharge passage was better than that with the standard 11-vane diffuser. At the low volume flows when the standard 11-vane diffuser gave better performance. The modified scroll gave performance characteristics that were practically the same as those of the standard scroll except at high tip speeds, where the peak performance was improved.

Description: The sound field of a rotating propeller is teated theoretically on the basis of aerodynamic principles. For the lower harmonics, the directional characteristics and the radiated sound energy are determined and are in conformity with existing experimental results.

Description: On the basis of the investigations so far completed on the behavior of PTL power plants under various operating conditions, in which the influence of the propeller characteristics is of considerable importance, the most important aspects of a control system for turbine-propeller jet power plants are deduced. A simple possible means for its concrete realization, which is also applicable to TL [NACA comment: TL, jet] power plants, is presented by means of examples. A control device of this kind is now being developed.

Description: A theoretical analysis of the temperature distribution through the trailing portion of a blade near the coolant passages of liquid cooled gas turbines was made. The analysis was applied to obtain the hot spot temperatures at the trailing edge and influence of design variables. The effective gas temperature was varied from 2000 degrees to 5000 degrees F in each investigation.

Description: Axial blowers are gaining importance as aircraft engine superchargers. However, the pressure head obtainable per stage is small. Due to the necessary great number of stages, the physical length of the blower becomes too great for an airworthy device. This report discusses several types of construction that permit a reduction in the length of the blower.

Description: An altitude-wind-tunnel investigation of a TG-100A gas turbine-propeller engine was performed. Pressure and temperature data were obtained at altitudes from 5000 to 35000 feet, compressor inlet ram-pressure ratios from 1.00 to 1.17, and engine speeds from 800 to 13000 rpm. The effect of engine speed, shaft horsepower, and compressor-inlet ram-pressure ratio on pressure and temperature distribution at each measuring station are presented graphically.

Description: A wind tunnel investigation was conducted to determine the performance of a 4000-pound-thrust axial-flow turbojet engine with a high flow compressor. Pressure altitudes included 5000 to 40000 feet with ram pressure ratios from 1.00 to 1.82. Altitudes included 20000 to 40000 feet and ram pressure ratios from 1.09 to 1.75. A comparison is made between engine performance with high flow and low flow compressors.

Description: A wind tunnel investigation was conducted to determine the performance of a turbine operating as an integral part of a turbojet engine. Data was obtained while the engine was running over full operable range of speeds at various altitudes and flight mach numbers, and with four nozzles of different outlet areas.A maximum turbine efficiency of 0.875 was obtained at altitude of 15 thousand feet, Mach number 0.53, and corrected turbine speed of 5900 rpm.

Description: The preignition characteristics of the R-2800 cylinder, as effected by fuel consumption, engine operating variables, and spark plug type and condition, were evaluated. The effects on preignition-limited performance of various percentages of aromatics (benzene, toluene, cumene, xylene) in a base fuel of triptane were investigated. Two paraffins (triptane and S + 6.0 ml TEL/gal) and two refinery blends (28-R and 33-R) were preignition rated. The effect of changes in the following engine operating variables on preignition limit was determined: inlet-air temperature, rear spark plug gasket temperature, engine speed, spark advance, tappet clearance, and oil consumption. Preignition limits of the R-2800 cylinder using Champion C34S and C35S and AC-LS86, LS87, and LS88 spark plugs were established and the effect of spark plug deterioration was investigated. No definite trends in preignition-limited indicated mean effective pressure were indicated for aromatics as a class when increased percentages of different aromatics were added to a base fuel of triptane. Three types of fuel (aromatics, paraffins, and refinery blends) showed a preignition range for this cylinder from 65 to 104 percent when based on the performance of S plus 6.0 ml TEL per gallon as 100 percent. The R-2800 cylinder is therefore relatively insensitive to fuel composition when compared to a CFR F-4 engine, which had a pre-ignition range from 72 to 100 percent for the same fuels. Six engine operating variables were investigated with the following results: preignition-limited indicated mean effective pressure decreased, with increases in engine speed, rear spark plug gasket temperature, inlet-air temperature, and spark advance beyond 20 F B.T.C. and was unaffected by rate of oil consumption or by tappet clearance. Spark plugs were rated over a range of preignition-limited indicated mean effective pressure from 200 to 390 pounds per square inch at a fuel-air ratio of 0.07 in the following order of increased resistance to preignition: AC-LS97, AC-LS88, Champion C358, AC-LS86, and Champion C34S. Spark plug deterioration in the form of cracks in the porcelain had been broken away from the center electrode and were retained in the spark plug cavity, the preignition limit was decreased as much as 57 percent. When the broken pieces had been removed, the preignition limit increased from that of the undamaged porcelain as the weight of removed porcelain was increased.

Description: Rim cracking in turbine wheels with welded blades was evaluated. The problem is explained on the basis of the occurrence of plastic flow in the rim during transient starting conditions when thermal compressive stresses resulting from high-temperature gradients exceed the proportional elastic limit of the material.

Description: Temperature and pressure distributions for an original and modified 3000 pound thrust axial flow turbojet engine were investigated. Data are included for a range of simulated altitudes from 5000 to 45000 feet, Mach numbers from 0.24 to 1.08, and corrected engine speeds from 10,550 to 13,359 rpm.

Description: A Douglas C-74 airplane, during a test dive at about 0.525 Mach number, experienced uncontrollable longitudinal oscillations sufficient to cause shedding of the outer wing panels and the subsequent crash of the airplane. Tests of a section of the horizontal tail plane from a C-74 airplane were conducted in the Ames 16-foot high-speed wind tunnel to investigate the possibility of the tail as a contributing factor to the accident. The results of the investigations of fabric-covered elevators in various conditions of surface deformation are presented in this report.

Description: Attempts were made to alleviate the buffeting of external fuel tanks mounted under the wings of a twin-engine Navy fighter plane. The Mach number at which the buffeting began was increased from 0.529 to 0.640 by streamlining the sway braces and increasing the lateral rigidity of the sway brace system. Further increases of the Mach number, at which buffeting began to 0.725, was obtained by moving the external fuel tank to a position under the fuselage.

Description: This report contains the flight-test results of the lateral and directional-stability and control phase (including tests with wing-tip tanks) of a general flying-qualities investigation of the Lockheed P-80A airplane (Army No. 44-85099). These tests were conducted at indicated airspeeds up to 494 miles per hour (0.691 Mach number) at low altitude and up to 378 miles per hour (0.816 Mach number) at high altitude. These tests showed that the flying qualities of the airplane were for the most part in accordance with the requirements of the Army Air Forces Stability and Control Specifications. The only major deficiency noted was the negative lateral stability with the wing-tip tanks installed.

Description: An altitude-wind-tunnel investigation has been made to determine the performance of a Curtiss 732-1C2-0 four-blade propeller on a YP-47M airplane at high blade loadings and engine power. Propeller characteristics were obtained for a range of power coefficients from 0.30 to 1.00 at free-stream Mach numbers of 0.40 and .50.

Description: A simulated altitude performance of a 25 1/2-inch-diameter annular-type turbojet combustor was performed to determine the effect of the distribution of basket-hole area on the altitude operational limits of the engine as imposed by the combustor.Total pressure drop was recorded, as well as the effect of fuel-nozzle flow capacity,and fuel-nozzle spray angle for one basket configuration. General observations were made for all configurations regarding flames, extent of afterburning, and durability of the baskets.

Description: An investigation was conducted to evaluate the operational characteristics of a 3000 pound thrust axial flow turbojet engine over a range of simulated altitudes from 2000 to 50,000 feet and simulated flight Mach numbers from 0 to 1.04 throughout the operable range of engine speeds. Engine operating range, acceleration, deceleration, starting, altitude, and flight Mach number compensation of the fuel control system, and operation of the lubrication system at high and low ambient air temperatures were evaluated.

Description: Partial measurements in flight of the handling qualities of a Grumman XF7F-1 airplane were investigated. Results are for low altitude and normal center of gravity conditions only.

Description: An investigation was conducted in the Cleveland altitude wind tunnel to determine the performance of a Curtiss propeller with four 838-lC2-lSRl blades on a YP-47M airplane at high blade loadings and engine powers. The study was made for a range of power coefficients between 0.30 and 1.00 at free-stream Mach numbers of 0.40 and 0.50. The results of the force measurements indicate primarily the trend of propeller efficiency for changes in power coefficient or advance-diameter ratio, inasmuch as corrections for the effects of tunnel-wall constriction on the installation have not been applied. Slip-stream pressure surveys across the propeller disk are presented to illustrate blade thrust load distribution for several operating conditions. At a free-stream Mach number of 0.40, nearly constant peak efficiencies were obtained at power coefficients from 0.30 to 0.70. A change in power coefficient from 0.70 to 0.90 reduced the peak efficiency about 5 percent. Blade stall at the tip sections became evident for a power coefficient of 0.91 when the advance-diameter ratio was reduced to 1.87. At a free-stream Mach number of 0.50, the highest propeller efficiencies were obtained for power coefficients from 0.80 to 1.00 at advance-diameter ratios above 2.90. At advance-diameter ratios below 2.90, the highest efficiencies were obtained for power coefficients of 0.60 and 0.70. The envelope of the efficiency curves decreased about 12 percent between advance-diameter ratios of 2.60 and 4.20. Local compressibility effects became evident for a power coefficient of 0.40 when the advance-diameter ratio was decreased to 1.75.

Description: As part of an investigation of the performance and operational characteristics of the axial-flow gas turbine-propeller engine, conducted in the Cleveland altitude wind tunnel, the performance characteristics of the compressor and the turbine were obtained. The data presented were obtained at a compressor-inlet ram-pressure ratio of 1.00 for altitudes from 5000 to 35,000 feet, engine speeds from 8000 to 13,000 rpm, and turbine-inlet temperatures from 1400 to 2100 R. The highest compressor pressure ratio obtained was 6.15 at a corrected air flow of 23.7 pounds per second and a corrected turbine-inlet temperature of 2475 R. Peak adiabatic compressor efficiencies of about 77 percent were obtained near the value of corrected air flow corresponding to a corrected engine speed of 13,000 rpm. This maximum efficiency may be somewhat low, however, because of dirt accumulations on the compressor blades. A maximum adiabatic turbine efficiency of 81.5 percent was obtained at rated engine speed for all altitudes and turbine-inlet temperatures investigated.

Description: Performance properties and operational characteristics of an axial-flow gas turbine-propeller engine were determined. Data are presented for a range of simulated altitudes from 5,000 to 35,0000 feet, compressor inlet- ram pressure ratios from 1.00 to 1.17, and engine speeds from 8000 to 13,000 rpm.

Description: An investigation was conducted to determine the operational and performance characteristics of the TG-100A gas turbine-propeller engine II. Windmilling characteristics were deterined for a range of altitudes from 5000 to 35,000 feet, true airspeeds from 100 to 273 miles per hour, and propeller blade angles from 4 degrees to 46 degrees.

Description: The NACA 23012-4 airfoil was investigated for the purpose of increasing lift by means of blowing out air from the wing, in conjunction with the effect of plain flap of variable contour and slotted flap of 25 percent chord length. The wing also was provided with a hinged nose, to be deflected at will. Air was blown out frcm the wing immediately in front of the flap; also at the opening between wing and hinged nose,tangentially to the surface of the wing. Another device employed to increase maximum lift was a movable slat, to be opened to form a clot. Lift was measured in relation to the volume of blown-out air and considerable increases were observed with increasing volume.

Description: A preliminary investigation of an axial-flow gas turbine-propeller engine was conduxted. Performance data were obtained for engine speeds from 8000 to 13,000 rpm and altitudes from 5000 to 35,000 feet and compressor inlet ram pressure ratios from 1.00 to 1.17.

Description: A 19XB-1 combustor was operated under conditions simulating zero-ram operation of the 19XB-1 turbojet engine at various altitudes and engine speeds. The combustion efficiencies and the altitude operational limits were determined; data were also obtained on the character of the combustion, the pressure drop through the combustor, and the combustor-outlet temperature and velocity profiles. At altitudes about 10,000 feet below the operational limits, the flames were yellow and steady and the temperature rise through the combustor increased with fuel-air ratio throughout the range of fuel-air ratios investigated. At altitudes near the operational limits, the flames were blue and flickering and the combustor was sluggish in its response to changes in fuel flow. At these high altitudes, the temperature rise through the combustor increased very slowly as the fuel flow was increased and attained a maximum at a fuel-air ratio much leaner than the over-all stoichiometric; further increases in fuel flow resulted in decreased values of combustor temperature rise and increased resonance until a rich-limit blow-out occurred. The approximate operational ceiling of the engine as determined by the combustor, using AN-F-28, Amendment-3, fuel, was 30,400 feet at a simulated engine speed of 7500 rpm and increased as the engine speed was increased. At an engine speed of 16,000 rpm, the operational ceiling was approximately 48,000 feet. Throughout the range of simulated altitudes and engine speeds investigated, the combustion efficiency increased with increasing engine speed and with decreasing altitude. The combustion efficiency varied from over 99 percent at operating conditions simulating high engine speed and low altitude operation to less than 50 percent at conditions simulating operation at altitudes near the operational limits. The isothermal total pressure drop through the combustor was 1.82 times as great as the inlet dynamic pressure. As expected from theoretical considerations, a straight-line correlation was obtained when the ratio of the combustor total pressure drop to the combustor-inlet dynamic pressure was plotted as a function of the ratio of the combustor-inlet air density to the combustor-outlet gas density. The combustor-outlet temperature profiles were, in general, more uniform for runs in which the temperature rise was low and the combustion efficiency was high. Inspection of the combustor basket after 36 hours of operation showed very little deterioration and no appreciable carbon deposits.

Description: The sea-level performance of I-16 turbojet engine at zero ram was investigated to determine the effects of an intake duct, shroud, and tail pipe intended for installation in an XFR-1 airplane. Engine speeds ranged from 8000 to 16,500 rpm for several variations of the intake duct and tail pipes.

Description: An investigation of a heated jet was conducted in conjunction with tests of an axial-flow jet-propulsion engine in the Cleveland altitude wind tunnel. Pressure and temperature surveys were made across the jet 10 and 15 feet behind the jet-nozzle outlet of the engine. Surveys were obtained at pressure altitudes of 10,000, 20,000, 30,000, and 40,000 feet with test-section velocities from 30 to 110 feet per second and test-section temperatures from 60 F to -50 F. From measurements taken throughout the operable range of engine speeds, tail-pipe outlet temperatures from 500 F to 1250 F and jet velocities from 400 to 2200 feet per second were obtained. The jet-survey data presented extend the work previously done with low-velocity and low-temperature jets to the region of high velocities and high temperatures. The results obtained agree with previously determined experimental data and with predicted theoretical expressions for the dimensionless transverse velocity and temperature profiles across a jet. The spread of both the temperature and the velocity profiles was very nearly linear. Dimensionless plots of temperature and velocity along the axis of a heated jet agree with experimental results of tests with a cold jet.

Description: This paper discussed the theory and design of dynamic "pressure augmentors" (diaphragms equal orifice plates and nozzles) and various forms of "pressure multipliers" (simple venturi tubes, Rateau-type multiple venturis, and a combination of shaped nozzle and simple venturi developed by the author). No complete theory of pressure multiplication is yet available; conditions of governing are discussed in relation to pressure-augmenting devices fitted either on the suction or the pressure side of the blower; fluctuations of output and power consumption caused by the presence of an augmentor are analyzed with the result that fitting on the pressure side appears generally preferable. Some considerations on the suitable design and selection of pressure-augmenting devices are appended.

Description: Flight and ground investigations have been made to compare an exhaust-ejector installation with a standard exhaust-collector-ring installation on air-cooled aircraft engines in a twin-engine airplane. The ground investigation allowed that, whereas the standard engine would have overheated above 600 horsepower, the engine with exhaust ejectors cooled at take-off operating conditions at zero ram. The exhaust ejectors provided as much cooling with cowl flaps closed as the conventional cowl flaps induced when full open at low airspeeds. The propulsive thrust of the exhaust-ejector installation was calculated to be slightly less than the thrust of the collector-ring-installation.

Description: The performance of a mixed-flow impeller in combination with a semivaneless diffuser were experimentally investigated. The diameter of the impeller was 11.0 inches and a maximum tip diameter of 14.74 inches. The semivaneless diffuser had an overall diameter of 28.00 inches. The performance properties of the mixed-flow impeller were also investigated with a 34.00 inch vane loss diffuser having a transition section of the same geometry as the semivaneless diffuser.

Description: The calculation of infinitesimal conical supersonic flow has been applied first to the simplest examples that have also been calculated in another way. Except for the discovery of a miscalculation in an older report, there was found the expected conformity. The new method of calculation is limited more definitely to the conical case.

Description: Operating characteristics of the 11-stage 4000-pound-thrust axial-flow turbojet engine were determined. A standard compressor and a compressor with the blade angles of the rotor and stator blades increased 5 degrees to obtain greater air flow, were investigated.

Description: The Russian AM 35 and AM 38 aircraft engines have superchargers with a swirl throttle, which appears to be a purely Russian development. This paper gives the results of test runs of the two engines, including the effects of the swirl throttle on engine performance.

Description: An altitude-wind-tunnel investigation has been made to determine the performance of Hamilton Standard 6507A-2 four-blade and three-blade propellers on a YP-47M airplane at high blade loadings and high engine powers. Characteristics of the four-blase propeller were obtained for a range of power coefficients from 0.10 to 1.00 at free-stream Mach numbers of 0.20, 0.30, 0.40. Characteristics of the three-blade propeller were obtained for a range of power coefficients from 0.30 to 1.00 at a free-stream Mach number of 0.40. Results of the force measurements indicate primarily the trend of propeller efficiency for changes in power coefficient or advance-diameter ratio because no corrections for the effects of tunnel-wall constriction on the installation were applied. Slipstream surveys are presented to illustrate blade thrust load distribution for certain operating conditions. Within the range of advance-diameter ratios investigated at each free-stream Mach number, the efficiency of the four-blade propeller decreased as the power coefficient was increased from 0.10 to 1.00. For the three-blade propeller, nearly constant maximum efficiencies were obtained for power coefficients from 0.32 to 0.63 at advance-diameter ratios between 1.90 and 3.00. In general, for conditions below the stall and critical tip Mach number, the maximum thrust load shifted from the inboard sections toward the tip sections as the power coefficient was increased or as the advance-diameter ratio was decreased. For conditions beyond the stall or critical tip Mach number, losses in thrust occurred on the outboard blade sections owing to flow break-down; the thrust load increased slightly on the inboard sections.

Description: A study was made of heat transfer in turbine blades and the effects on blade temperature of cooling the blade root and tip, changing the dimensions of the blades, raising the cycle temperatures, insulating with ceramics, and cooling by circulation of air or water through hollow blades.

Description: Combustion chamber performance properties of a 3000-pound-thrust axial-flow turbojet engine were determined. Data are presented for a range of simulated altitudes from 15,000 to 45,0000 feet and a range of Mach numbers from 0.23 to 1.05 for various modifications of the engine.